1. Field of the Invention
The present invention relates to a gas turbine compressor that can improve the efficiency of the compressor by maintaining optimal clearance between the moving blades and the rear case rings of the compressor during operation. The present invention further relates to a clearance controlling method for a gas turbine compressor.
2. Description of Related Art
In a gas turbine plant, compressed air from a gas turbine compressor is guided into a combustor, and the high-temperature gas generated when the compressed air is combusted together with a fuel is guided into the gas turbine to drive the gas turbine. A typical design is one in which a portion of the compressed air is bled and directed to the stationary and moving blades of the gas turbine side and is used to cool these blades.
FIG. 3 is a cross-sectional view showing the typical structure of the connecting member between the gas turbine compressor and the gas turbine in a conventional gas turbine plant as described above. As shown in FIG. 3, gas turbine compressor 1 is provided with a plurality of stationary blades 3 which are fixed in place inside case rings 2a in case 2, and a plurality of moving blades 5 which are attached to the periphery of each disk 4 which are coaxially fixed on the rotor side (rotor not shown). These stationary blades 3 and moving blades 5 are disposed so as to alternate with one another along the shaft orientation of the rotor and are designed to compress and send compressed air in the direction indicated by arrow f1 through rotation of the rotor.
The compressed air sent from gas turbine compressor 1 is directed to combustor 6 in which the compressed air is mixed with fuel and combusted to form a combustion gas. As it expands, the combustion gas hg passes through stationary blades 8 in gas turbine 7 and rotates moving blades 9, thereby rotationally driving the rotor (not shown) in gas turbine 7. At the same time, a portion (4xcx9c10% of the main flow, for example) of the compressed air which is flowing inside gas turbine compressor 1 becomes bleed air f2, and is taken up inside bleeding chamber 2b which is formed inside case 2. After being expelled out from case 2 via a flange 2c which is provided so as to communicate with bleeding chamber 2b, bleed air f2 is guided to stationary blades 8 and moving blades 9 on the gas turbine 7 side and cools these blades.
Bleeding chamber 2b is a ring-shaped space formed between each case ring 2a and case main body 2d which covers over the periphery of these case rings 2a. In the axial direction of the rotor of gas turbine 1, bleeding chamber 2 is provided so as to be overlapped the area where the end surfaces of respective case rings 2a face one another. In other words, the space intervals between facing end surfaces of case rings 2a form bleeding holes 2e for bleeding air into bleeding chamber 2b from compressed air f1, which is the main flow. Bleed air f2 is guided into bleeding chamber 2b toward the radial direction of the rotor with passing through bleeding holes 2e. Bleed air f2 is then quickly expelled outside via flange 2c for cooling.
Clearance of specific dimensions is maintained between moving blades 5 and case rings 2a in order to avoid contact between them during operation in gas turbine compressor 1. The clearance varies during operation depending on difference in thermal expansion between case 2, and disks 4 and moving blades 5. If the difference in thermal expansion between these parts becomes too large, the efficiency of the compressor in gas turbine compressor 1 may remarkably deteriorate. In view of this problem, it is necessary to adopt optimal clearance dimensions in the design after taking this factor into consideration.
In fact, however, thermal deformation of case rings 2a is complicated, so that it is difficult to provide a design that fits these circumstances with excellent precision. Namely, in compressor rear case ring 2a1 among case rings 2a, which is positioned at a later stage (i.e., most downstream position) in the compressor and in particular, has a great influence on the efficiency of the compressor in gas turbine compressor 1, the temperature of the main flow (compressed air f1) flowing inside compressor rear case ring 2a1 remarkably increases, for example, from 370xc2x0 C. to 460xc2x0 C., due to an increase in enthalpy by the effect of compression.
As a result of such large temperature difference, the overall shape of compressor rear case ring 2a when undergoing thermal expansion will deform such that it becomes gradually wider toward the direction of flow of compressed air f1. Accordingly, the clearance formed between moving blades 5 and the inner surface of compressor rear case rings 2a becomes gradually wider toward downstream from upstream, and therefore, the clearance does not have uniform dimensions.
Accordingly, the clearance dimensions during operation are not uniform along the axial direction of compressor rear case rings 2a1. This makes it difficult to provide a design that ensures optimal clearance dimensions, therefore, it becomes difficult to improve the efficiency of the compressor in gas turbine compressor 1.
The present invention was conceived in view of the above-described circumstances and has an object that is to provide a gas turbine compressor that can improve the efficiency of the compressor by maintaining optimal dimensions for the clearance formed between the ends of the moving blades and the inner surface of the rear case rings of the compressor during operation. The present invention has the other object that is to provide a clearance controlling method for a gas turbine compressor.
The present invention employs the following means to resolve the above-described problems.
A gas turbine compressor according to a first aspect of the present invention comprises a plurality of moving blades which are provided around rotor disks and rotate together with the rotor disks; compressor rear case rings surrounding the periphery of these moving blades and forming a compression flow path therein; and a bleeding chamber which is provided around the compressor rear case rings and introduces a portion of a main flow moving through the compression flow path as bleed air; and a cooling flow path which is formed between the compressor rear case rings and the bleeding chamber in which bleed air on its way to the bleeding chamber flows along the outer surface of the compressor rear case rings.
According to the above-described gas turbine compressor, the main flow moving inside the compressor rear case rings is dependent on a compressing effect such that the temperature increases as the main flow moves toward downstream. As a result, the compressor rear case rings are heated from the inside. However, the bleed air which flows through the cooling flow path cools the compressor rear case rings from their periphery, so that temperature gradient becomes small along the axial direction. As a result, when the compressor rear case rings undergo thermal expansion, the expansion in the direction of their diameters at each position along the axial direction is roughly equivalent, making it easy to predict the behavior of thermal deformation. Accordingly, a design for ensuring optimal clearance is easily provided. Thus, the dimensions of the clearance that is formed between the ends of the moving blades and the inner surface of the compressor rear case rings can be optimized, so that the efficiency of the compressor can be further improved.
Furthermore, in the above-described gas turbine compressor, the cooling flow path may comprise boundaries in the axial direction when viewed in a cross-section that includes the axis of the compressor rear case rings, in which the boundaries include at least a region extending from a position on the upstream edge of the outer surface of the compressor rear case rings to a position at a furthest downstream corresponding to the moving blades.
According to the above gas turbine compressor, in the regions along the axial direction of the compressor rear case rings, the boundaries required for controlling clearance in particular are securely cooled. The dimensions of the clearance formed between the ends of the moving blades and the inner surface of the compressor rear case rings during operation can be maintained at optimal values with certainty. Accordingly, the efficiency of the compressor can be improved.
In the above-described gas turbine compressor, a sleeve in the shape of a ring or an interrupted ring is disposed so as to cover the bleeding flow intake for bleeding a portion of the main flow moving in the compression flow path. This bleed air may be made to flow along the outer surface of the compressor rear case rings.
According to this gas turbine compressor, the compressor case rings can be cooled from their periphery with certainty.
Moreover, in this gas turbine compressor, the shape of the cooling flow path when viewed upstream of the main flow may be scallop-shaped.
According to this gas turbine compressor, the heating surface area of the cooling flow path is larger than that of when a simple curved surface shape is used, therefore, higher cooling effects can be obtained.
In this gas turbine compressor, the compressor rear case rings may be made of a material having low linear expansion.
According to this gas turbine compressor, the thermal expansion can be reduced even at the same metal temperature, so that the overall amount of thermal expansion of the compressor rear case rings can be reduced. As a result, the clearance that is formed between the ends of the moving blades and the inner surface of the compressor rear case rings can be held to a small value. Accordingly, the efficiency of the compressor can be further improved.
Furthermore, in this gas turbine, a heat shield coating may be applied to the inner surface of the compressor rear casing rings.
According to this gas turbine compressor, the amount of heat introduced from the main flow can be reduced as a result of the heat shield coating for the same boundary conditions (i.e., for the same main flow temperature conditions). As a result, the metal temperature of the compressor rear case rings can be reduced, and the amount of thermal expansion overall can be reduced. Accordingly, the clearance formed between the ends of the moving blades and the inner surface of the compressor rear case rings can be held to a small value. Accordingly, the efficiency of the compressor can be further improved.
A method for controlling clearance in a gas turbine compressor according to a second aspect of the present invention is a method for controlling a clearance formed between ends of moving blades and an inner surface of compressor rear case rings in a gas turbine compressor which comprises a plurality of moving blades which are provided around rotor disks and rotate together with the rotor disks; compressor rear case rings surrounding the periphery of these moving blades and forming a compression flow path therein; and a bleeding chamber which is provided around the compressor rear case rings and introduces a portion of a main flow moving through the compression flow path as bleed air; the method comprising the steps of flowing bleed air on its way to the bleeding chamber along an outer surface of the compressor rear case rings, and introducing the bleed air into the bleeding chamber.
According to the above-described clearance controlling method for a gas turbine compressor, the main flow moving in the compressor rear case rings is subjected to a compressing effect so that the temperature increases as the flow moves further downstream. As a result, the compressor rear case rings are heated from the inside. However, the bleed air that flows over the outer surface cools the compressor rear case rings from their periphery, so that the temperature gradient becomes small along the axial direction. As a result, when the compressor rear case rings undergo thermal expansion, the expansion in the direction of their diameters at each position along the axial direction is roughly equivalent, making it easy to predict the behavior of thermal deformation. Accordingly, a design for ensuring optimal clearance is easily provided. Thus, the dimensions of the clearance that is formed between the ends of the moving blades and the inner surface of the compressor rear case rings can be optimized, so that the efficiency of the compressor can be further improved.
Furthermore, in the above-described clearance controlling method for a gas turbine compressor, flow boundaries of the bleed air to the outer surface when viewed in a cross-section that includes the axis of the compressor rear case rings, in which the boundaries at least the region extending from a position on the upstream edge of the outer surface to a furthest downstream position corresponding to the moving blades.
According to this clearance controlling method for a gas turbine compressor, the boundaries required for controlling clearance in particular are securely cooled along the axial direction of the compressor rear case rings. The dimensions of the clearance formed between the ends of the moving blades and the inner surface of the compressor rear case rings during operation can be optimized confidently. Accordingly, the efficiency of the compressor can be improved.